Flexible support structure for a geared architecture gas turbine engine

ABSTRACT

A gas turbine engine according to an example of the present disclosure includes, among other things, a fan shaft driving a fan having fan blades, a fan shaft support that supports the fan shaft and a gear system connected to the fan shaft. The gear system includes a ring gear defining a ring gear transverse stiffness, a gear mesh defining a gear mesh transverse stiffness, and a reduction ratio greater than 2.3. The ring gear transverse stiffness is less than 20% of the gear mesh transverse stiffness.

CROSS REFERENCE TO RELATED APPLICATIONS

The present disclosure is a continuation of U.S. patent application Ser.No. 14/859,381, filed Sep. 21, 2015, which is a continuation of U.S.patent application Ser. No. 14/604,811, filed Jan. 26, 2015, now U.S.Pat. No. 9,239,012, issued Jan. 19, 2016, which is acontinuation-in-part of U.S. patent application Ser. No. 13/623,309,filed Sep. 20, 2012, now U.S. Pat. No. 9,133,729, issued Sep. 15, 2015,which is a continuation-in-part of U.S. application Ser. No. 13/342,508,filed Jan. 3, 2012, now U.S. Pat. No. 8,297,916, issued Oct. 30, 2012,which claimed priority to U.S. Provisional Application No. 61/494,453,filed Jun. 8, 2011.

BACKGROUND

The present disclosure relates to a gas turbine engine, and moreparticularly to a flexible support structure for a geared architecturetherefor.

Epicyclic gearboxes with planetary or star gear trains may be used ingas turbine engines for their compact designs and efficient high gearreduction capabilities. Planetary and star gear trains generally includethree gear train elements: a central sun gear, an outer ring gear withinternal gear teeth, and a plurality of planet gears supported by aplanet carrier between and in meshed engagement with both the sun gearand the ring gear. The gear train elements share a common longitudinalcentral axis, about which at least two rotate. An advantage of epicyclicgear trains is that a rotary input can be connected to any one of thethree elements. One of the other two elements is then held stationarywith respect to the other two to permit the third to serve as an output.

In gas turbine engine applications, where a speed reduction transmissionis required, the central sun gear generally receives rotary input fromthe powerplant, the outer ring gear is generally held stationary and theplanet gear carrier rotates in the same direction as the sun gear toprovide torque output at a reduced rotational speed. In star geartrains, the planet carrier is held stationary and the output shaft isdriven by the ring gear in a direction opposite that of the sun gear.

During flight, light weight structural cases deflect with aero andmaneuver loads causing significant amounts of transverse deflectioncommonly known as backbone bending of the engine. This deflection maycause the individual sun or planet gear's axis of rotation to loseparallelism with the central axis. This deflection may result in somemisalignment at gear train journal bearings and at the gear teeth mesh,which may lead to efficiency losses from the misalignment and potentialreduced life from increases in the concentrated stresses.

SUMMARY

A gas turbine engine according to an example of the present disclosureincludes a fan shaft configured to drive a fan, a support configured tosupport at least a portion of the fan shaft, the support defining asupport transverse stiffness and a support lateral stiffness, a gearsystem coupled to the fan shaft, and a flexible support configured to atleast partially support the gear system. The flexible support defines aflexible support transverse stiffness with respect to the supporttransverse stiffness and a flexible support lateral stiffness withrespect to the support lateral stiffness. The input defines an inputtransverse stiffness with respect to the support transverse stiffnessand an input lateral stiffness with respect to the support lateralstiffness.

In a further embodiment of any of the forgoing embodiments, the supportand the flexible support are mounted to a static structure.

In a further embodiment of any of the forgoing embodiments, the staticstructure is a front center body of the gas turbine engine.

In a further embodiment of any of the forgoing embodiments, the flexiblesupport is mounted to a planet carrier of the gear system, and the inputis mounted to a sun gear of the gear system.

In a further embodiment of any of the forgoing embodiments, the fanshaft is mounted to a ring gear of the gear system.

In a further embodiment of any of the forgoing embodiments, the gearsystem is a star system.

In a further embodiment of any of the forgoing embodiments, the flexiblesupport is mounted to a ring gear of the gear system, and the input ismounted to a sun gear of the gear system.

In a further embodiment of any of the forgoing embodiments, the fanshaft is mounted to a planet carrier of the gear system.

In a further embodiment of any of the forgoing embodiments, the flexiblesupport transverse stiffness and the input transverse stiffness are bothless than the support transverse stiffness.

In a further embodiment of any of the forgoing embodiments, the flexiblesupport transverse stiffness and the input transverse stiffness are eachless than about 20% of the support transverse stiffness.

In a further embodiment of any of the forgoing embodiments, the flexiblesupport transverse stiffness and the input transverse stiffness are eachless than about 11% of the support transverse stiffness.

In a further embodiment of any of the forgoing embodiments, the input tothe gear system is coupled to a turbine section, and the gear system isconfigured to drive a compressor rotor at a common speed with the fanshaft.

A gas turbine engine according to an example of the present disclosureincludes a fan shaft configured to drive a fan, a support configured tosupport at least a portion of the fan shaft, and a gear systemconfigured to drive the fan shaft. The gear system includes a gear meshthat defines a gear mesh transverse stiffness and a gear mesh lateralstiffness. A flexible support is configured to at least partiallysupport the gear system. The flexible support defines a flexible supporttransverse stiffness with respect to the gear mesh transverse stiffnessand a flexible support lateral stiffness with respect to the gear meshlateral stiffness. The input defines an input transverse stiffness withrespect to the gear mesh transverse stiffness and an input lateralstiffness with respect to the gear mesh lateral stiffness.

In a further embodiment of any of the forgoing embodiments, both theflexible support transverse stiffness and the input transverse stiffnessare less than the gear mesh transverse stiffness.

In a further embodiment of any of the forgoing embodiments, the flexiblesupport transverse stiffness is less than about 8% of the gear meshtransverse stiffness, the input transverse stiffness is less than about5% of the gear mesh transverse stiffness, and a transverse stiffness ofa ring gear of the gear system is less than about 20% of the gear meshtransverse stiffness.

In a further embodiment of any of the forgoing embodiments, a transversestiffness of a planet journal bearing which supports a planet gear ofthe gear system is less than or equal to the gear mesh transversestiffness.

In a further embodiment of any of the forgoing embodiments, the supportand the flexible support are mounted to a front center body of the gasturbine engine.

A method of designing a gas turbine engine according to an example ofthe present disclosure includes providing a fan shaft, and providing asupport configured to support at least a portion of the fan shaft, thesupport defining at least one of a support transverse stiffness and asupport lateral stiffness, and providing a gear system coupled to thefan shaft. The gear system includes a gear mesh that defines a gear meshlateral stiffness and a gear mesh transverse stiffness. The methodincludes providing a flexible support configured to at least partiallysupport the gear system, and providing an input to the gear system. Theflexible support defines a flexible support transverse stiffness withrespect to the gear mesh transverse stiffness and a flexible supportlateral stiffness with respect to the gear mesh lateral stiffness. Theinput defines an input transverse stiffness with respect to the gearmesh transverse stiffness and an input lateral stiffness with respect tothe gear mesh lateral stiffness.

In a further embodiment of any of the forgoing embodiments, the flexiblesupport lateral stiffness is less than the gear mesh lateral stiffness,and the flexible support transverse stiffness is less than the gear meshtransverse stiffness.

In a further embodiment of any of the forgoing embodiments, both theflexible support transverse stiffness and the input transverse stiffnessare less than the gear mesh transverse stiffness.

The various features and advantages of this invention will becomeapparent to those skilled in the art from the following detaileddescription of an embodiment. The drawings that accompany the detaileddescription can be briefly described as follows.

BRIEF DESCRIPTION OF THE DRAWINGS

Various features will become apparent to those skilled in the art fromthe following detailed description of the disclosed non-limitingembodiment. The drawings that accompany the detailed description can bebriefly described as follows:

FIG. 1 is a schematic cross-section of a gas turbine engine;

FIG. 2 is an enlarged cross-section of a section of the gas turbineengine which illustrates a fan drive gear system (FDGS);

FIG. 3 is a schematic view of a flex mount arrangement for onenon-limiting embodiment of the FDGS;

FIG. 4 is a schematic view of a flex mount arrangement for anothernon-limiting embodiment of the FDGS;

FIG. 5 is a schematic view of a flex mount arrangement for anothernon-limiting embodiment of a star system FDGS;

FIG. 6 is a schematic view of a flex mount arrangement for anothernon-limiting embodiment of a planetary system FDGS.

FIG. 7 is a schematic view of a flex mount arrangement for anothernon-limiting embodiment of a star system FDGS;

FIG. 8 is a schematic view of a flex mount arrangement for anothernon-limiting embodiment of a planetary system FDGS;

FIG. 9 shows another embodiment; and

FIG. 10 shows yet another embodiment.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct defined within a nacelle 15, while the compressor section 24drives air along a core flow path C for compression and communicationinto the combustor section 26 then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with two-spoolturbofans as the teachings may be applied to other types of turbineengines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a first (or low) pressure compressor 44 and afirst (or low) pressure turbine 46. The inner shaft 40 is connected tothe fan 42 through a speed change mechanism, which in exemplary gasturbine engine 20 is illustrated as a geared architecture 48 to drivethe fan 42 at a lower speed than the low speed spool 30. The high speedspool 32 includes an outer shaft 50 that interconnects a second (orhigh) pressure compressor 52 and a second (or high) pressure turbine 54.A combustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 57 of the engine static structure 36 is arranged generally betweenthe high pressure turbine 54 and the low pressure turbine 46. Themid-turbine frame 57 further supports bearing systems 38 in the turbinesection 28. The inner shaft 40 and the outer shaft 50 are concentric androtate via bearing systems 38 about the engine central longitudinal axisA which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, and fan section 22 may be positioned forward or aftof the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present invention isapplicable to other gas turbine engines including direct driveturbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, withthe engine at its best fuel consumption—also known as “bucket cruiseThrust Specific Fuel Consumption (‘TSFC’)”—is the industry standardparameter of lbm of fuel being burned divided by lbf of thrust theengine produces at that minimum point. “Low fan pressure ratio” is thepressure ratio across the fan blade alone, without a Fan Exit Guide Vane(“FEGV”) system. The low fan pressure ratio as disclosed hereinaccording to one non-limiting embodiment is less than about 1.45. “Lowcorrected fan tip speed” is the actual fan tip speed in ft/sec dividedby an industry standard temperature correction of [(Tram ° R)/(518.7°R)]^(0.5). The “Low corrected fan tip speed” as disclosed hereinaccording to one non-limiting embodiment is less than about 1150ft/second.

With reference to FIG. 2, the geared architecture 48 generally includesa fan drive gear system (FDGS) 60 driven by the low speed spool 30(illustrated schematically) through an input 62. The input 62, which maybe in the form of a coupling, both transfers torque from the low speedspool 30 to the geared architecture 48 and facilitates the segregationof vibrations and other transients therebetween. In the disclosednon-limiting embodiment, the FDGS 60 may include an epicyclic gearsystem which may be, for example, a star system or a planet system.

The input coupling 62 may include an interface spline 64 joined, by agear spline 66, to a sun gear 68 of the FDGS 60. The sun gear 68 is inmeshed engagement with multiple planet gears 70, of which theillustrated planet gear 70 is representative. Each planet gear 70 isrotatably mounted in a planet carrier 72 by a respective planet journalbearing 75. Rotary motion of the sun gear 68 urges each planet gear 70to rotate about a respective longitudinal axis P.

Each planet gear 70 is also in meshed engagement with rotating ring gear74 that is mechanically connected to a fan shaft 76. Since the planetgears 70 mesh with both the rotating ring gear 74 as well as therotating sun gear 68, the planet gears 70 rotate about their own axes todrive the ring gear 74 to rotate about engine axis A. The rotation ofthe ring gear 74 is conveyed to the fan 42 (FIG. 1) through the fanshaft 76 to thereby drive the fan 42 at a lower speed than the low speedspool 30. It should be understood that the described geared architecture48 is but a single non-limiting embodiment and that various other gearedarchitectures will alternatively benefit herefrom.

With reference to FIG. 3, a flexible support 78 supports the planetcarrier 72 to at least partially support the FDGS 60A with respect tothe static structure 36 such as a front center body which facilitatesthe segregation of vibrations and other transients therebetween. Itshould be understood that various gas turbine engine case structures mayalternatively or additionally provide the static structure and flexiblesupport 78. It should be understood that lateral as defined herein isgenerally transverse to the axis of rotation A and the term “transverse”refers to a pivotal bending movement with respect to the axis ofrotation A which typically absorbs deflection applied to the FDGS 60.The static structure 36 may further include a number 1 and 1.5 bearingsupport static structure 82 which is commonly referred to as a “K-frame”which supports the number 1 and number 1.5 bearing systems 38A, 38B.Notably, the K-frame bearing support defines a lateral stiffness(represented as Kframe in FIG. 3) and a transverse stiffness(represented as Kframe^(BEND) in FIG. 3) as the referenced factors inthis non-limiting embodiment.

In this disclosed non-limiting embodiment, the lateral stiffness (KFS;KIC) of both the flexible support 78 and the input coupling 62 are eachless than about 11% of the lateral stiffness (Kframe). That is, thelateral stiffness of the entire FDGS 60 is controlled by this lateralstiffness relationship. Alternatively, or in addition to thisrelationship, the transverse stiffness of both the flexible support 78and the input coupling 62 are each less than about 11% of the transversestiffness (Kframe^(BEND)). That is, the transverse stiffness of theentire FDGS 60 is controlled by this transverse stiffness relationship.

With reference to FIG. 4, another non-limiting embodiment of a FDGS 60Bincludes a flexible support 78′ that supports a rotationally fixed ringgear 74′. The fan shaft 76′ is driven by the planet carrier 72′ in theschematically illustrated planet system which otherwise generallyfollows the star system architecture of FIG. 3.

With reference to FIG. 5, the lateral stiffness relationship within aFDGS 60C itself (for a star system architecture) is schematicallyrepresented. The lateral stiffness (KIC) of an input coupling 62, alateral stiffness (KFS) of a flexible support 78, a lateral stiffness(KRG) of a ring gear 74 and a lateral stiffness (KJB) of a planetjournal bearing 75 are controlled with respect to a lateral stiffness(KGM) of a gear mesh within the FDGS 60.

In the disclosed non-limiting embodiment, the stiffness (KGM) may bedefined by the gear mesh between the sun gear 68 and the multiple planetgears 70. The lateral stiffness

(KGM) within the FDGS 60 is the referenced factor and the staticstructure 82′ rigidly supports the fan shaft 76. That is, the fan shaft76 is supported upon bearing systems 38A, 38B which are essentiallyrigidly supported by the static structure 82′. The lateral stiffness(KJB) may be mechanically defined by, for example, the stiffness withinthe planet journal bearing 75 and the lateral stiffness (KRG) of thering gear 74 may be mechanically defined by, for example, the geometryof the ring gear wings 74L, 74R (FIG. 2).

In the disclosed non-limiting embodiment, the lateral stiffness (KRG) ofthe ring gear 74 is less than about 12% of the lateral stiffness (KGM)of the gear mesh; the lateral stiffness (KFS) of the flexible support 78is less than about 8% of the lateral stiffness (KGM) of the gear mesh;the lateral stiffness (KJB) of the planet journal bearing 75 is lessthan or equal to the lateral stiffness (KGM) of the gear mesh; and thelateral stiffness (KIC) of an input coupling 62 is less than about 5% ofthe lateral stiffness (KGM) of the gear mesh.

With reference to FIG. 6, another non-limiting embodiment of a lateralstiffness relationship within a FDGS 60D itself are schematicallyillustrated for a planetary gear system architecture, which otherwisegenerally follows the star system architecture of FIG. 5.

It should be understood that combinations of the above lateral stiffnessrelationships may be utilized as well. The lateral stiffness of each ofstructural components may be readily measured as compared to filmstiffness and spline stiffness which may be relatively difficult todetermine.

By flex mounting to accommodate misalignment of the shafts under designloads, the FDGS design loads have been reduced by more than 17% whichreduces overall engine weight. The flex mount facilitates alignment toincrease system life and reliability. The lateral flexibility in theflexible support and input coupling allows the FDGS to essentially‘float’ with the fan shaft during maneuvers. This allows: (a) the torquetransmissions in the fan shaft, the input coupling and the flexiblesupport to remain constant during maneuvers; (b) maneuver inducedlateral loads in the fan shaft (which may otherwise potentially misaligngears and damage teeth) to be mainly reacted to through the number 1 and1.5 bearing support K-frame; and (c) both the flexible support and theinput coupling to transmit small amounts of lateral loads into the FDGS.The splines, gear tooth stiffness, journal bearings, and ring gearligaments are specifically designed to minimize gear tooth stressvariations during maneuvers. The other connections to the FDGS areflexible mounts (turbine coupling, case flex mount). These mount springrates have been determined from analysis and proven in rig and flighttesting to isolate the gears from engine maneuver loads. In addition,the planet journal bearing spring rate may also be controlled to supportsystem flexibility.

FIG. 7 is similar to FIG. 5 but shows the transverse stiffnessrelationships within the FDGS 60C (for a star system architecture). Thetransverse stiffness (KIC^(BEND)) of the input coupling 62, a transversestiffness (KFS^(BEND)) of the flexible support 78, a transversestiffness (KRG^(BEND)) of the ring gear 74 and a transverse stiffness(KJB^(BEND)) of the planet journal bearing 75 are controlled withrespect to a transverse stiffness (KGM^(BEND)) of the gear mesh withinthe FDGS 60.

In the disclosed non-limiting embodiment, the stiffness (KGM^(BEND)) maybe defined by the gear mesh between the sun gear 68 and the multipleplanet gears 70. The transverse stiffness (KGM^(BEND)) within the FDGS60 is the referenced factor and the static structure 82′ rigidlysupports the fan shaft 76. That is, the fan shaft 76 is supported uponbearing systems 38A, 38B which are essentially rigidly supported by thestatic structure 82′. The transverse stiffness (KJB^(BEND)) may bemechanically defined by, for example, the stiffness within the planetjournal bearing 75 and the transverse stiffness (KRG^(BEND)) of the ringgear 74 may be mechanically defined by, for example, the geometry of thering gear wings 74L, 74R (FIG. 2).

In the disclosed non-limiting embodiment, the transverse stiffness(KRG^(BEND)) of the ring gear 74 is less than about 12% of thetransverse stiffness (KGM^(BEND)) of the gear mesh; the transversestiffness (KFS^(BEND)) of the flexible support 78 is less than about 8%of the transverse stiffness (KGM^(BEND)) of the gear mesh; thetransverse stiffness (KJB^(BEND)) of the planet journal bearing 75 isless than or equal to the transverse stiffness (KGM^(BEND)) of the gearmesh; and the transverse stiffness (KIC^(BEND)) of an input coupling 62is less than about 5% of the transverse stiffness (KGM^(BEND)) of thegear mesh.

FIG. 8 is similar to FIG. 6 but shows the transverse stiffnessrelationship within the FDGS 60D for the planetary gear systemarchitecture.

FIG. 9 shows an embodiment 200, wherein there is a fan drive turbine 208driving a shaft 206 to in turn drive a fan rotor 202. A gear reduction204 may be positioned between the fan drive turbine 208 and the fanrotor 202. This gear reduction 204 may be structured, mounted andoperate like the gear reduction disclosed above. A compressor rotor 210is driven by an intermediate pressure turbine 212, and a second stagecompressor rotor 214 is driven by a turbine rotor 216. A combustionsection 218 is positioned intermediate the compressor rotor 214 and theturbine section 216.

FIG. 10 shows yet another embodiment 300 wherein a fan rotor 302 and afirst stage compressor 304 rotate at a common speed. The gear reduction306 (which may be structured, mounted and operate as disclosed above) isintermediate the compressor rotor 304 and a shaft 308, which is drivenby a low pressure turbine section.

It should be understood that relative positional terms such as“forward,” “aft,” “upper,” “lower,” “above,” “below,” and the like arewith reference to the normal operational attitude of the vehicle andshould not be considered otherwise limiting.

It should be understood that like reference numerals identifycorresponding or similar elements throughout the several drawings. Itshould also be understood that although a particular componentarrangement is disclosed in the illustrated embodiment, otherarrangements will benefit herefrom.

Although particular step sequences are shown, described, and claimed, itshould be understood that steps may be performed in any order, separatedor combined unless otherwise indicated and will still benefit from thepresent disclosure.

The foregoing description is exemplary rather than defined by thelimitations within. Various non-limiting embodiments are disclosedherein, however, one of ordinary skill in the art would recognize thatvarious modifications and variations in light of the above teachingswill fall within the scope of the appended claims. It is therefore to beunderstood that within the scope of the appended claims, the disclosuremay be practiced other than as specifically described. For that reasonthe appended claims should be studied to determine true scope andcontent.

What is claimed is:
 1. A gas turbine engine, comprising: a fan shaftdriving a fan having fan blades; a fan shaft support that supports saidfan shaft; a gear system connected to said fan shaft, said gear systemincludes a ring gear defining a ring gear transverse stiffness, a gearmesh defining a gear mesh transverse stiffness, and a reduction ratiogreater than 2.3; and wherein said ring gear transverse stiffness isless than 20% of said gear mesh transverse stiffness.
 2. The gas turbineengine of claim 1, further comprising a bypass ratio greater than ten(10).
 3. The gas turbine engine of claim 2, further comprising a lowpressure turbine with an inlet, an outlet, and a low pressure turbinepressure ratio greater than 5:1, wherein said low pressure turbinepressure ratio is a ratio of a pressure measured prior to said inlet asrelated to a pressure at said outlet prior to any exhaust nozzle.
 4. Thegas turbine engine of claim 3, further comprising a two stage highpressure turbine.
 5. The gas turbine engine of claim 2, furthercomprising a low corrected fan tip speed less than about 1150 ft/second,wherein said low corrected fan tip speed is an actual fan tip speed atan ambient temperature divided by [(Tram ° R)/(518.7° R)]0.5, where Trepresents said ambient temperature in degrees Rankine.
 6. The gasturbine engine of claim 2, further comprising a gear system input tosaid gear system defining a gear system input transverse stiffness andsaid fan shaft support defining a fan shaft support transversestiffness, wherein said gear system input transverse stiffness is lessthan 20% of said fan shaft support transverse stiffness.
 7. The gasturbine engine of claim 6, further comprising a mid-turbine frameincluding at least one airfoil extending into a flow path, and whereinsaid fan shaft support is a K-frame bearing support.
 8. The gas turbineengine of claim 6, wherein said gear mesh defines a gear mesh lateralstiffness and said gear system input defines a gear system input lateralstiffness and said gear system input lateral stiffness is less than 5%of said gear mesh lateral stiffness.
 9. The gas turbine engine of claim8, wherein said fan shaft support defines a fan shaft support lateralstiffness and said gear system input lateral stiffness is less than 11%of said fan shaft support lateral stiffness.
 10. The gas turbine engineof claim 6, further comprising a two stage high pressure turbine and alow fan pressure ratio of less than 1.45 and said low fan pressure ratiois measured across the fan blades alone.
 11. The gas turbine engine ofclaim 10, further comprising a low pressure turbine with an inlet, anoutlet, and a low pressure turbine pressure ratio greater than 5:1,wherein said low pressure turbine pressure ratio is a ratio of apressure measured prior to said inlet as related to a pressure at saidoutlet prior to any exhaust nozzle.
 12. The gas turbine engine of claim11, wherein said gear system input transverse stiffness is less than 11%of said fan shaft support transverse stiffness.
 13. The gas turbineengine of claim 11, further comprising a low corrected fan tip speedless than about 1150 ft/second, wherein said low corrected fan tip speedis an actual fan tip speed at an ambient temperature divided by [(Tram °R)/(518.7° R)]0.5, where T represents said ambient temperature indegrees Rankine.
 14. The gas turbine engine of claim 13, wherein saidgear system input transverse stiffness is less than 11% of said fanshaft support transverse stiffness, and said gear system input bothtransfers torque and facilitates segregation of vibrations.
 15. A gasturbine engine, comprising: a fan shaft driving a fan having fan blades;a fan shaft support that supports said fan shaft; a gear systemconnected to said fan shaft, said gear system includes a ring geardefining a ring gear transverse stiffness, a gear mesh defining a gearmesh transverse stiffness, and a reduction ratio greater than 2.3;wherein said ring gear transverse stiffness is less than 20% of saidgear mesh transverse stiffness; and a low fan pressure ratio of lessthan 1.45 and said low fan pressure ratio is measured across the fanblades alone.
 16. The gas turbine engine of claim 15, further comprisinga two stage high pressure turbine and a low pressure turbine with aninlet, an outlet, and a low pressure turbine pressure ratio greater than5:1, wherein said low pressure turbine pressure ratio is a ratio of apressure measured prior to said inlet as related to a pressure at saidoutlet prior to any exhaust nozzle.
 17. The gas turbine engine of claim16, further comprising a low corrected fan tip speed less than about1150 ft/second, wherein said low corrected fan tip speed is an actualfan tip speed at an ambient temperature divided by [(Tram ° R)/(518.7°R)]0.5, where T represents said ambient temperature in degrees Rankine.18. The gas turbine engine of claim 17, further comprising a mid-turbineframe including at least one airfoil extending into a flow path.
 19. Thegas turbine engine of claim 15, further comprising a low corrected fantip speed less than about 1150 ft/second, wherein said low corrected fantip speed is an actual fan tip speed at an ambient temperature dividedby [(Tram ° R)/(518.7° R)]0.5, where T represents said ambienttemperature in degrees Rankine.
 20. A gas turbine engine, comprising: afan shaft driving a fan having fan blades; a fan shaft support thatsupports said fan shaft; a gear system connected to said fan shaft, saidgear system includes a ring gear defining a ring gear transversestiffness and a ring gear lateral stiffness, a gear mesh defining a gearmesh transverse stiffness and a gear mesh lateral stiffness, and areduction ratio greater than 2.3; and wherein said ring gear transversestiffness is less than 20% of said gear mesh transverse stiffness andsaid ring gear lateral stiffness is less than 12% of said gear meshlateral stiffness.
 21. The gas turbine engine of claim 20, furthercomprising a low fan pressure ratio of less than 1.45 and said low fanpressure ratio is measured across the fan blades alone.
 22. The gasturbine engine of claim 21, wherein said ring gear transverse stiffnessis less than 12% of said gear mesh transverse stiffness.
 23. The gasturbine engine of claim 21, further comprising a low pressure turbinewith an inlet, an outlet, and a low pressure turbine pressure ratiogreater than 5:1, wherein said low pressure turbine pressure ratio is aratio of a pressure measured prior to said inlet as related to apressure at said outlet prior to any exhaust nozzle.
 24. The gas turbineengine of claim 23, further comprising a low corrected fan tip speedless than about 1150 ft/second, wherein said low corrected fan tip speedis an actual fan tip speed at an ambient temperature divided by [(Tram °R)/(518.7° R)]0.5, where T represents said ambient temperature indegrees Rankine.
 25. The gas turbine engine of claim 24, wherein saidring gear transverse stiffness is less than 12% of said gear meshtransverse stiffness.
 26. The gas turbine engine of claim 25, furthercomprising a mid-turbine frame including at least one airfoil extendinginto a flow path.
 27. The gas turbine engine of claim 21, furthercomprising a low corrected fan tip speed less than about 1150 ft/second,wherein said low corrected fan tip speed is an actual fan tip speed atan ambient temperature divided by [(Tram ° R)/(518.7° R)]0.5, where Trepresents said ambient temperature in degrees Rankine.
 28. The gasturbine engine of claim 20, further comprising a high speed spoolincluding a two stage high pressure turbine, and a low speed spoolincluding a low pressure turbine with an inlet, an outlet, and a lowpressure turbine pressure ratio greater than 5:1, wherein said lowpressure turbine pressure ratio is a ratio of a pressure measured priorto said inlet as related to a pressure at said outlet prior to anyexhaust nozzle.
 29. The gas turbine engine of claim 28, wherein saidgear system includes a planet carrier and said fan shaft is mounted tosaid planet carrier, and further comprising a low corrected fan tipspeed less than about 1150 ft/second, wherein said low corrected fan tipspeed is an actual fan tip speed at an ambient temperature divided by[(Tram ° R)/(518.7° R)]0.5, where T represents said ambient temperaturein degrees Rankine.
 30. The gas turbine engine of claim 29, furthercomprising at least one bearing system, and a three stage low pressurecompressor, wherein said high speed spool includes an outer shaft andsaid low speed spool includes an inner shaft, and said inner shaft andouter shaft are concentric and rotate in coordination with said at leastone bearing system about a longitudinal axis of said engine.